Spacecraft Power Systems

The Generation and Storage of Electrical Power D. B. Kanipe February 21, 2012

Power Systems Batteries Solar Cells + Batteries Fuel Cells

RTG Nuclear Reactors ?

The Power System controls the generation, storage, and efficient use of power

Must also provide protection against a failure in one circuit inducing a failure in other systems

2

Power System Design Drivers Customer/User requirements Mission, distance from the sun Spacecraft configuration

Mass constraints Dimensional constraints Launch Vehicle constraints Thermal constraints

Expected lifetime Attitude control system

Pointing Effectors Viewing

Orbit Payload requirements

Power type, voltage, current Duty cycle, peak load Fault protection

Mission constraints Maneuver rates G-loads

3

Power System Functional Block Diagram

4

Power Source

Source Control

Power Distribution, Main Bus Control & Main Bus Protection

Power Conditioning Load

Energy Storage

Energy Storage Control

Main Bus

- Batteries - Solar - RTG - Fuel Cells - Nuclear - R Dynamic - Solar Dynamic

- Shunt Regulator - Series Regulator - Shorting Switch Array

- DC-DC conversion - DC-AC conversion - Voltage regulator

- Battery charge control - Voltage regulator

Design Practice (1/2) Direct Current Switching

Switches or relays: positive line to an element with a direct connection to “ground’ on negative side Therefore, element is inert until commanded

Arc Suppression Locate as close to the source of the arc as possible Conductive cables, connectors, solar array edges, any

current-carrying element should not be exposed to the ambient plasma

Modularity Simplifies testing Easier element replacement Reduces “collateral” damage

5

Design Practice (2/2) Grounding

Cause of some debate among EEs Common ground cable preferable to individual component

grounding Easier to maintain a common potential Less likely to disturb sensitive components Can be difficult to do in large spacecraft

Sometimes it is necessary to completely isolate an element from other spacecraft noise

Continuity Avoid buildup of static potential; i.e., any voltage difference Any shielding must have continuity and a common ground

Complexity KISS

6

Batteries: Definition of Terms (1/2) Charge Capacity, Cchg

Total electric charge stored in a battery; measured in amp-hours (e.g., 40A for 1 hour = 40Ah)

Energy Capacity, Ebat Total energy stored in a battery; Cchg X Vavg (Joules or watt-hours)

Average Discharge Voltage, Vavg (Number of cells in series) X (Cell discharge voltage)

Depth of Discharge, DOD Percent of battery capacity used in discharge cycle 75% DOD means 25% remaining Try to limit DOD to promote longer cycle life

7

Batteries: Definition of Terms (2/2)

Charge Rate, Rchg Rate at which the battery can accept charge

(amps/unit time) Energy Density, ebat

Energy per unit mass stored in battery Joules/kg or Watt-hours/kg

Two categories of batteries

Primary Secondary

8

Primary Batteries Long storage capability (missile in a silo) Dry (without electrolyte) until needed

Activation by introducing electrolyte into dry battery Electrolyte may be solid at room temperature.

Activate heater to melt electrolyte. (Thermal battery) Used for early major mission event

Short duration May be isolated from major power bus Usually non-rechargeable Mass penalty

9

Secondary Batteries Typically, lower energy density, Rechargeable DOD management required

Will be in eclipse ~ 40% of orbit 12-16 discharge/charge cycles per day Battery degradation and lifetime reduction Given spacecraft power usage and maximum allowable

DOD for lifetime design: DOD = Energy required during eclipse Stored battery energy

DOD = 𝑃𝐿𝑡𝑑𝐶𝑐𝑐𝑐𝑉𝑎𝑎𝑐

= 𝑃𝐿𝑡𝑑𝐸𝑏𝑎𝑡

PL = load power, Watts

td = discharge time, hours

10

Recharge Rate Must manage charge rate Too fast heat and damage Too slow can’t make up discharge

Rule of thumb 𝑅𝑐𝑐𝑐 = 𝐶𝑐𝑐𝑐

15𝑐= 𝐼𝑐𝑐𝑐 Charging current

I.E., a battery can accept about 1/15 of its total capacity per hour (conservative)

“Trickle” charge = 𝐶𝑐𝑐𝑐45𝑐

11

DOD Management Typically, a LEO spacecraft spends 40% of its time

discharging and 60% charging DOD during eclipse is limited by the rate at which it can

be restored in sunlight All expended energy must be restored or you will have a

net drain Driver ends up being the charge rate

DOD limited to 7-8% per orbit Battery temperature can affect charge rate

Battery generally must be charged at a voltage > Vavg (~20% higher) to restore full charge

This affects solar array design

12

Solar Arrays Photoelectric Effect

Electrons are emitted from matter as a result of absorption of short wavelength electromagnetic radiation such as visible light.

Originally limited to skin acreage Deployables offer more flexibility, but also more

complexity Solar Cell Characteristics

1st order: V decreases as T increases (and vice versa) 2nd order: I increases as T increases

But only about 10% relative to the voltage drop Therefore, overall power output is reduced: as

temperature increases: P = I*V May need radiators to remove excess heat

13

Maximum Power Point (MPP) Desirable to operate at the MPP if

possible Minimize mass and maximize efficiency

14

I

V

MPP

Maximum area rectangle under the IV curve

Sun Tracking Best, of course, if sun is normal to

the array Cosine rule applies – up to a point

Optimum

Up to 45-60˚ cosine function works, then falls off rapidly

15

Beta and Alpha Beta, ß Angle between a line from the sun to the

center of the earth, and spacecraft orbit

ß

Alpha, α Apparent rotation of sun angle from

spacecraft pov during its orbit. α = 0-360

16

E

Solar to Electric Conversion Efficiency Gallium arsenide solar cells (Ga-As) More efficient (20%) and radiation tolerant More expensive

Crystalline Silicon Cells 11-16%, 18-20%, >20%?

Multi-junction (multi-layer cells) Top layer converts light in the visible range Bottom layer(s) optimized for infrared Up to 30% efficiency Not surprisingly, expensive

17

Radioisotope Thermoelectric Generator (RTG)

Renders spacecraft independent of the sun Converts heat energy generated by radioisotope decay

into direct current energy via thermoelectric effect Plutonium 238, 238Pu Strontium 90, 90Sr

Expensive Complicated ground handling

Heat Radiation

RTG radiation detrimental to spacecraft electronics

18

Fuel Cells Direct conversion of chemical energy

into electricity Like batteries, but more efficient Oxidizer and fuel fed into a cell Electricity generated from oxidation

reaction in the cell Space applications use oxygen/hydrogen By product: water ~ 35% efficiency

19

Power Conditioning & Control Voltage from power source, especially

solar arrays, may fluctuate Power conditioning functions Control solar array output Control battery charge/discharge cycle Regulate voltage supplied to spacecraft

systems Two types of systems Dissipative (direct energy transfer, DET) Non-dissipative (peek power tracking, PPT)

20

Dissipative Systems Simpler Not in series with array output

Dissipates current in excess of instantaneous load requirement

21

Solar Array

Shunt Battery Charge

Controller

Spacecraft Loads

Battery

Non-dissipative Systems (PPT) In series regulation of solar power

22

Usually reserved for large spacecraft

Solar Array

Battery Charge

Controller

Spacecraft Loads

Battery

Battery Discharge Controller

PPT

Additional Power Sources Nuclear Reactors Dynamic Isotope Systems Alkali Metal Thermal-to-Electric

Conversion (AMTEC) Solar Dynamic

23